Full report

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'. \ , 'I, NASA TECHNICAL MEMORANDUM '\ NASA TM-77050 TRANSONIC CRYOGENIC TEST SECTION FOR THE GBTTINGEN TUBE FACILITY H. Hornung, G. Hefer, P. Krogmann and E. Stanewsky NASA-TM-77050 19840008151 Translation of "Transsonische Kryomesstrecke fUr den GBttinger Rohrwindkana1," Deutsche Forschungs- und Versuchsansta1t fUr Luft- und Raumfahrt (DFVLR), Aerodynamische Versuchsansta1t GBttingen, GBttingen, w. Germany, Report IB 222 - 82 A 19, May 3, 1982, pp. 1-19 LIBRARY COpy. LANGlt y qr""RcH CET&1l J , ! ':'( , lYtJ4 LIBRARY. NASA HAMPTON, VGlt1'1 Kt=: VI s ED' VE es I orJ NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D.C. 20546 /" DECEMBER 1983 l 1111111111111~~~~~!IIII~11\1111

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.. . -12. STANDARD TITLE PAGE 1. R,orl Ho. GonTluunl Att... ion Ho• 3: R.t/pllnl·.-COIO/O; No. 'NASA TM-77050 4. TII/. end SuL'ill. • s. RIDa;' Dol. TRANSONIC CRYOGENIC TEST SECION FOR DECEMBER 1983 THE GTTINGEN TUBE FAGILITY 7. ""60,(,) H. Hornung, G. Heier"P, Krogmann, and E. Stanewsky 9. Pnlo,lII:ng O';OIi lellen HCIII. Q'ld A.dd,u, ·Leo Kanner Associates Redwood City, California 94063 12. }ponsefln; A.glnt)' Hern. ond Add,u. •• Pllrormln; O,gonilolion Cod. .. P•.ror.lnll O,gcniZGlion R.po,' No. 10. Work UnU No. JJ. Conrroct or C,on I No. NASW-354l 13. T". 01 Rlport end Peried Cour.tt Translation ~ National Aeronautics and Spa ce Adminis- tic. Sponaorln; A;onc)' Cad. tration, Washington, D.C. 20546 15. Su;;:h:::rnlcry NOlu Translation Qf "Transsonisthe Kryomesstrecke fUr den Gttinger Rohrwindkanal," Deutsche Forschungs- und .Versuchsanstalt fUr Luft- und Raumfahrt (DFVLR), Aerodynamiscoe Versuchsanstalt Gattingen, Gttingen, W. ermany, Repor B 222 - 82 J6. Asl,cicl A l~, May 3, 1982, pp. 1-19. The design of modern aircraft requires the solution of proolems related to transonic· flow at high Reynolds numbers. To investigate these problems expeTimentally, it is proposed to. extend the· Ludwieg·tube facility in Gttingen by adding a transonic cryogenic test ection. After stating the rquirements for such a tst ection, the technrca1 concept _is briefly explained and a preliminary estimate-pf the costs is given. t;l :' .." . 17. K.y h,d. (Selectees tl)' AuthOr(S» 18. Dlllrlbulion SIoIIClanl .. -Upclassified --Unlimited -. 19. s.tU"')' ClolIl'. (of ,hi. ,;p,,,I) 20. SIC'lrTII), Cloul'. (0' Ihls pog.) Unclassified Unclassified .0 ., ...'- 21. Ho. of Pog.' I 22. . . -.:tJ (~3 '7 7 0 - ! - .0 ... . -. . I : N 8 3- ;;t-3 3;;;- fj?-=It-

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TABLE OF CONTENTS 1. Introduction Page 1 2. Requirements on a Research Tunnel 1 2.1. Reynolds Number Range 2.2. Main Dimensions 3. Technical Concept 4. Costs 5. Summary ii 2 4 5 6 7

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LIST OF ABBREVIATIONS NTF National Transonic Facility DFVLR Deutsche Forschung- und Versuchsanstalt fUr Luftund Raumfahrt (German Aerospace Research Establishment) ETW European Transonic Wind Tunnel iii

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LIST OF TABLES AND FIGURES Table 1. Problem Area and Requirements on the Research Tunnel Table 2. t'lork and Cos t Development Plan Fig. 1. Reynolds Number Dependency of the Lift and Regions of Different Wind Tunnels. Airfoil CAST 10-2/DOA 2. Ma = 0.765; a = 2°. Fig. 2. Aerodynamic Performance Parameter (cA/c ) ·Ma. W Measurements in the 0.3 Meter TCT of NASA, Langley. Fig. 3. Mach Number, Reynolds Number, Range of Some DFVLR Wind Tunnels. Fig. 4. Technical Concept. Fig. 5. Location Diagram, Scale 1:500. iv

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TRANSONIC CRYOGENIC TEST SECTION FOR THE G5TTINGEN TUBE FACILITY H. Hornung, G. Hefer, P. Krogmann and E. Stanewsky Institute for Experimental Fluid Mechanics in GBttingen 1. Introduction /5* Economy and productivity of modern aircraft are highly dependent upon the solution of aerodynamic problems that can be collectively arranged under the heading "Transonic Flow Phenomena at High Reynolds Numbers." The necessity of constructing wind tunnels necessary to solve these problems has been recognized for quite some time by all countries involved in the science of flight and aircraft construction and is manifested in the construction of the National Transonic Facility (NTF) in the USA and the planning of the European Transonic Wind Tunnel (ETW). As the design data of both large tunnels intended for 6 aircraft development -- Re = 120.10 for NTF and Re x = max rna 6 = 50'10 for ETW -- show, there are different opinions concerning the Reynolds number range to be covered. 2. Requirements on a Research Tunnel Within the DFVLR Institute for Experimental Fluid Mechanics is working on a solution of partial problems from the complex area of transonics. The facilities TKG, TWB and HKG, which belong to the wind tunnel division, are currently at our disposal to work on these problems experimentally. These facilities are unsuitable for longer term research programs with expanded experimental investigations because of cost considerations -this also applies to the ETW. However, a more severe drawback is that they do not fulfill the necessary requirements for successful research in the areas cited, because the Reynolds numbers attainable are too low. For this reason, several suggestions for constructing a transonic cryogenic wind tunnel in the DFVLR have been made in recent years, e.g. [1]. *Numbers in the margin indicate pagination in the foreign text. 1

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·. In addition to research problems, the plans foresaw making series measurements on total models. This increases the dimensions and expenditure of the apparatus and raises costs. -- The tunnel we propose is supposed to primarily serve the /6 purpose of examining physical flow phenomena, allowing lower construction costs. -- In this regard different wind tunnel principles were examined with special emphasis on their utility. The result was that a tube tunnel is best suited for fulfilling the requirements briefly mentioned below. It is therefore proposed to expand the GBttingen tube facilities accordingly. Recent advances in the design of transport aircraft were obtained by the use of transonic airfoils. One can assume that the continued development of such airfoils, e.g. in conjunction with an active boundary layer control, will also in the future allow considerable improvements in flight performance. For that reason the tunnel should be designed in such a way that airfoils and the flow phenomena appearing upon them can be investigated under realistic conditions. This requirement dominates to a great degree the design of the tunnel. 2.1. Reynolds Number Range The flow around transonic airfoils in all flow ranges can be highly dependent upon the Reynolds number [2]. This is represented in Figs. 1 and 2 by the example of lift and the aerodynamic performance parameter in Breguet's cruising distance equation, (cA/c ) ·Ma. In addition to the great Reynolds W 6 number dependency in the whole investigated range of Re = 2.10 6 to Re = 45.10 , it is noteworthy that the curves give no clue whatsoever regarding the behavior of the aerodynamic parameters 6 at the Reynolds numbers of Re > 45·10 .* 6 *Re = 45.10 is the maximum attainable Reynolds number in the [footnote cont'd next page] 2

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, The results leave the question unanswered whether the exploitation of transonic potential during aircraft development requires wind tunnel experiments at flight Reynolds numbers, or if it is 17 possible to achieve this by extrapolating the results, beginning with an unknown upper Reynolds number limit. The latter would require knowledge of the flow phenomena into the range of the flight Reynolds number. In conjunction with the results mentioned above, it can be deduced: 6 -- that Reynolds numbers of Re > 45.10 must be achieved to investigate the flow phenomena that determine the Reynolds number sensitivity of a configuration. For a research tunnel with its low dimensions this would mean that it must be operated cryogenically. If one applies a minimum airfoild depth of t. = m~n = 150 mm, we would like to mention that the maximum attainable 6 Reynolds number in this tunnel will be Re = 70.10 • This covers a Reynolds number range that in many cases includes the flight Reynolds number (see Fig. 3). In such a wind tunnel, investigations on components can be conducted where the basic flow phenomena of a total configuration occur. This is demonstrated by the example of an aircraft with wings of high aspect ratio, which also determines the geometry of the tunnel: -- The flow around a wing of high aspect ratio depends largely upon the geometry of the basic airfoil. Therefore, the Reynolds number behavior of the configuration can first be studied on the basic airfoil. Flow phenomena that depend primarily on the Reynolds number and that also determine the Reynolds number sensitivity of an airfoil are, for example, footnote from p. 2, cont'd: the shock boundary layer 0.3 m transonic cryogenic tunnel of the NASA facility in Langley. The measurements were conducted as a cooperative effort of the DFVLR/AVA. See bibliographic reference [3]. 3

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interaction and the behavior of the boundary layer when exposed to sustained rear adverse pressure gradients. These phenomena, including separation effects, can be investigated on basic models attached to the lower wall of the test section at Reynolds numbers exceeding the ones in regular airfoil tests. -- In the future, investigations with respect to boundary 18 layer control will be paid much more attention. These tests can be conducted on airfoil and basic models at realistic Reynolds numbers. The tests planned in the research tunnel are briefly compiled in Table 1. This table also contains more, primarily secondary requirements. 2.2. Main Dimensions We are assuming that the minimum dimensions of the test section are determined by the requirements for airfoil measurements. The geometry determined in this manner will also suffice for the other requirements in Table 1. For reasons of technical measurement resolution, manufacturing accuracy and surface quality, and especially with regard to experiments concerning boundary layer control and the related machining of slots, we consider a minimum airfoil chord of t = 150 mm necessary for testing airfoils. Based on experimental values for the ratio of tunnel height to airfoil chord HIt = 3 and tunnel width to airfoil chord of BIt = 2, a minimum chord of tmin = 150 mm leads to a test section cross section of B x H = 300 x 450 mm. The naturally good flow quality in a Ludwieg tube is somewhat reduced by the boundary layer in the storage tube, the thickness of which increases with the tube Mach number. By 4

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using a nozzle with a contraction ratio of K > 3 the Mach number in the storage tube remains below MaR = 0.2; this guarantees a good flow quality. Thus, one obtains a tube diameter of approximately DR = 0.8 m. The requirement for a measurement time of about one second demands a tube length of LR = 130 m. 3 the downstream tank to be V ~ 2.V ~ 130 m , Assuming the volume of K R the main dimensions of the tunnel are established. 3. Technical Concept /9 The technical concept is represented in Fig. 4. It differs from the supersonic tunnels in GBttingen primarily by the arrangement of the fast acting valve which is located behind the transonic test section and serves simultaneously as a diffuser for adjusting the Mach number. With the exception of the tank, which can be insulated on the inside, all parts are constructed of cryogenically suitable material. The liquid nitrogen is stored in a tank provided by the supplier. Part of the gas retrieved in the tank during a run can be pumped back into the tube by means of a compressor, and the cold losses can be replaced by using liquid nitrogen. The most important performance data of the wind tunnel are compiled in Fig. 4. In addition to the model size, the Mach number and the stagnation temperature, the maximum Reynolds number is determined by the highest possible stagnation pressure. The stagnation pressure was limited to a value of Pomax = 10 bar according to an estimation of model deformations. If one assumes that the stagnation temperature is chosen in such a way that at a local Mach number of Ma = 1.4 saturation of M the nitrogen is just reached, then the upper curve of the number diagram depicted in Fig. 3 is I Reynolds number Mach obtained. When comparing this curve with those of the ETW and KKK it should be noted that for these tunnels, generally used for testing complete models, a Reynolds number reference length of a.l./s was utilized. K 5

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As a testing site the available tube facility of the Research Center in Gattingen offers special advantages, as many aggregates and measurement devices can be shared. Figure 5 shows a floor plan. It depicts the most advantageous solution from an operating standpoint, but it requires constructing an annex between houses 30 and 19, which is much preferred to maintaining the tunnel in house 30. 4. Costs /10 The following cost estimate for commercially available parts such as pipe, tank, etc., is based upon suggested retail price offers of relevant firms. Part costs that could only be calculated with the help of detailed plans are estimated by using older plans as a scale. The final investment sum will amount to 2.6 million German Marks. (Approximately $1,083,333.00. One US dollar equals 2.40 West German Marks, March, 1983). Cost Itemizin9:, TDM (thousands of marks) Pipe (including insulation) 320 Gate valve Nozzle Test section Fast-closing valve Tank Control, safety devices Construction Engineering expenses Total investment costs 350 100 350 450 100 280 450 200 2,600 TDM Table 2 is a labor and cost development plan. The operating expenses of the wind tunnel depend greatly upon the portion of cryogenic operating time, as primarily the 6

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expenses for liquid nitrogen become noticeable. These expenses are estimated at OM 150,000 (approximately $62,500) per annum. The expenditures can be reduced considerably by installing a refrigerator for recooling the used nitrogen. 5. Summary The economy and productivity of modern aircraft are highly dependent upon the solution of aerodynamic problems in transonic flow at high Reynolds numbers. Since basic problems are /11 currently largely unsolved, intensive research of these flow phenomena is necessary. As a test conducted in the summer of 1981 by the OFVLR in cooperation with NASA in the 0.3-m cryogenic wind tunnel at Langley demonstrated, transcribing experimental results from wind tunnels with too low Reynolds numbers to flight conditions is still not assured. For that reason, an experimental facility allowing basic examinations at flight Reynolds numbers at reasonable cost is necessary. At current levels of knowledge, this can only be economically feasible with cryogenic technology. The principle of the tube tunnel guarantees that the investigation can be conducted at the best possible flow quality. With the planned facility the OFVLR is creating a unique research tool for testing the most important current aerodynamic problems of flight technology, the solution of which is required for sensible industrial utilization of the ETW. 7

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REFERENCES 1. Ludwieg, H. and P. Krogmann, "The transonic cryogenic wind tunnel project of the German Aerospace Research Establishment," DFVLR Report IB 251-78 A 39, GBttingen, West Germany, 1978. 2. Stanewsky, E., Wechselwirkung zwischen der reibungsfreien AussenstrBmung und der Grenzschicht an transsonischen Profilen [Interaction between the outer inviscid flow and the boundary layer on transonic airfoils], Thesis, Technical University of Berlin, D83, 1981. 3. Ray, E., C. Johnson and E. Stanewsky, NASA/DFVLR Advanced Technology Airfoil Progrrun, Tests in the 0.3 meter cryogenic tunnel with the airfoil CAST 10-2/DOA 2, August, 1981. To be published as a NASA report. 8

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", TABLE 1. PROBLEM AREA AND REQUIREMENTS ON THE RESEARCH 113 TUNNEL. Problem Area 1. Wings with high aspect ratio - Airfoil tes ts - Thrust boundary layer Requirements Test section cross section: 0.3 x 0.45 m2 Schlieren window, floor trailing edge interference mountable model - Boundary layer interaction (removal by suction, heat transfer) Probe drive Auxiliary mechanisms 2. Wings with small aspect ratio Half model technology 3. Missiles - Influence of Mach and Reynolds number upon asymmetric vortex shedding 4. Flow quality 5. Wall interference 6. Measuring technique - Boundary layer and flow field measurements (average) - Measurement of unsteady phenomena - Development of short duration measurement methods Rectangular cross section Drive mechanisms for flow field measurements The investigation of its influence requires high flow quality. Therefore: contraction ratio tubel test section cross section 3:1 Exchangeable wind tunnel walls. Boundary layer suction on the side walls Measurement time 1 s Storage tube length 130 m 9

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I-' o TABLE 2. tvORK AND COST DEVELOPMENT PLAN /14 Hork Phase I----r--I-r-r-r-nIIT'I "I I ' 1. rotal concept ana !l.nal specl.il.cdt.iuJI Jctailed design proposals, issuance of 1st Year 2nd Year -3rd Year I I I I I I l--! I II ' I I I l' 1 I I 'Ii i- 3. wV~~ders.f- i I I Inr: i I Cnffioonent manufacture 3. 1 Storaqe tube and tank 3. 2 Test ection with nozzle, diffuser and startl.n] valve 3.3 LI2 supply and exhaust system 0) '04 ContIol system 4. ':onstruction L 1 '·oundation and alteratwn worK .1. 2 Storaqe tube and tank 4. 3 L -Svstem 2 • 4.-1 Test sectl.on· 4.5 Insulation 4.G Extension of data acqul.sition system 15. Functional testing, alterations, improvements Investent (thousndR nf ermn Mrs) . ! I I' I I I I ! . r--r-t-t-t--t--H~'IIH-l I_I I ! I I I I I I llT I I I I I I I I I I I I r---: JOO I. JOO 1.000 - _I

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/15 t---- TWB ----l t--------03MTCT, 2) -' OJ7S -I CA Free transition O}O OJ6S q60 qss ;>/ qso ~ 6 5 10 CFWTl) ,,/ CD/ / ,,/ Forced transition o TKG 6 CFWT o 0,3 MTCT 7 -' 5 10 Re 1) Lockheed Compressible Flow Wind Tunnel - - - - 2) 0.3 ~eter Iransonlc fryogenlc Tunnel NASJ-Langley Fig. 1. Reynolds number dependency of the lift and regions of different wind tunnels. Airfoil CAST 10-2/DOA 2. Ma = 0.765; a = 2°. 11

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I-' N 40 ~. c \ ~'Ma , C w ' ... 30 v/ / 1.( - 0 Ma =O,30 6 I Y-:1 .v )4/./ ..'-" / F ;.r 1/ vI' /' / I I Ma =O}65 Ci =2° Half darkened symbols: Forced transition 20 Open symbols: T I I J 6 5 10 Free transition I T -. r , 7 5 108 10 Rg Fig. 2. Aerodynamic performance parameter (cA/c ) ·Ma. Measurements in the 0.3 meter TCT of NASA, Langley. W I

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120 '105 I - - - ------------- /17 C'1ilrac ter Nr. model CD Total 100 @ model G) Components Re © (e.g. al.rfOl.lsl ® ® I 80 c,o ~<:- I 60 ",-" I Q)HDK 40 20 ®TKG o o Q5 l.:3 tl.C ~eference l.!ngth for Re O,l·VS 3· H K I ~.c ~ .V <:-"Y ? 8747 I 1,0 1,5 Ma Fig. 3. Mach number, Reynolds number range of some DFVLR wind tunnels. 13

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~ I-' .:::.. Tube Gate Valve Nozzle --- ---s: ~- 8 LN2 -' Test Section Fast Acting Valve Tanks -- ( Fig. 4. Technical concept. Tube Diameter 800 rom Length 130 m Nominal pressure 16 bar Gate Valve Cross section 0.3 x 0.45 Model depth 0.15 m 130 m3 Tank Volume Nominal pressure 6 bar Characteristic data: Maximum stagnation pressure 10 bar Temperature range 90-300 K 2 m Mach number range 0.2-1.2 Maximum Reynolds number 70.10 6 Measurement time 0.7-1.1 s ~

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\ \ House 21 \ \ ~.:Jew tube Present tubes -;-. \ \ \ \ \ House 30 Fig. 5. Location diagram, scale 1:500 . • f-' Ul I' It;

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